Bleed air systems for use with aircrafts and related methods

ABSTRACT

Bleed air systems for use with aircrafts and related methods are disclosed. An example apparatus includes a turbo-compressor including a compressor has a compressor inlet fluidly coupled to a low-pressure compressor of an aircraft engine and an intermediate port of a high-pressure compressor of the aircraft engine. The compressor inlet to receive fluid from either the low-pressure compressor or the high-pressure compressor based on a first system parameter of the aircraft. A turbine has a turbine inlet fluidly coupled to the intermediate port of the high pressure compressor and a high-pressure port of the high pressure compressor of the aircraft engine. The turbine inlet to receive fluid from either the intermediate port of the high-pressure compressor or the high-pressure port of the high-pressure compressor based on a second system parameter of the aircraft.

RELATED APPLICATION

This patent arises from a continuation-in-part of U.S. application Ser.No. 13/357,293 which was filed on Jan. 24, 2012, entitled “Bleed AirSystems for use with Aircrafts and Related Methods” and is herebyincorporated herein by reference in its entirety.

FIELD OF THE DISCLOSURE

The present disclosure relates generally to aircrafts and, moreparticularly, to bleed air systems for use with aircrafts and relatedmethods.

BACKGROUND

Commercial aircrafts or jetliners typically employ an environmentalcontrol system to pressurize a passenger cabin of the aircraft and/orthermal anti-icing systems to provide heated air for anti-icingapplications. Air supply to these systems is typically provided by bleedair extracted from or provided by a compressor of an aircraft engine. Tomeet pressure and/or temperature demands of the various aircraftsystems, bleed air is often extracted from a high stage of alow-pressure compressor of the aircraft engine. For example, bleed airis often extracted from an eighth stage compressor of an aircraftengine. The pressurized bleed air is then often cooled via a precoolerprior to providing the bleed air to a system of the aircraft (e.g.,environmental control system). Thus, much of the energy spent by theengine to produce the bleed air is wasted when cooling the bleed air viathe precooler. As a result, high-pressure bleed air extracted from thecompressor may significantly reduce the efficiency of the engine.

To reduce extraction of bleed air, some known systems employ aturbo-compressor that receives ambient air from an atmospheric inlet.The turbo-compressor pressurizes the ambient air prior to supplying thevarious aircraft systems. However, the atmospheric inlet produces drag.Additionally, the atmospheric inlet is often susceptible to icing and,thus, requires an anti-icing system that increases costs and systemcomplexity. Further, the compressor may have to be relatively large toproduce a pressure change sufficient to power the systems of anaircraft.

SUMMARY

An example apparatus includes a turbo-compressor. The turbo-compressorincludes a compressor has a compressor inlet fluidly coupled to alow-pressure compressor of an aircraft engine and an intermediate portof a high-pressure compressor of the aircraft engine. The compressorinlet to receive fluid from either the low-pressure compressor or thehigh-pressure compressor based on a first system parameter of theaircraft. A turbine has a turbine inlet fluidly coupled to theintermediate port of the high pressure compressor and a high-pressureport of the high pressure compressor of the aircraft engine. The turbineinlet to receive fluid from either the intermediate port of thehigh-pressure compressor or the high-pressure port of the high-pressurecompressor based on a second system parameter of the aircraft.

Another example apparatus includes a turbo-compressor having acompressor and a turbine. A first inlet passageway to fluidly couple alow-pressure port from an aircraft engine to a compressor inlet of thecompressor. A second inlet passageway to fluidly couple a firstintermediate port from the aircraft engine to the compressor inlet. Athird inlet passageway to fluidly couple a high-pressure port from anaircraft engine to a turbine inlet of the turbine. A fourth inletpassageway to fluidly couple a second intermediate port from theaircraft engine to the turbine inlet.

An example method includes fluidly coupling a compressor inlet of aturbo-compressor to a low-pressure bleed air source provided by alow-pressure compressor of an aircraft engine via a first inletpassageway and fluidly coupling the compressor inlet to an intermediatebleed air source provided by a high-pressure compressor of the aircraftengine via a second inlet passageway. The method includes fluidlycoupling a turbine inlet of the turbo-compressor to a high-pressurebleed air source provided by the high-pressure compressor of theaircraft engine via a third inlet passageway and fluidly coupling theturbine inlet and the intermediate bleed air source provided byhigh-pressure compressor via a fourth inlet passageway.

The features, functions and advantages that have been discussed can beachieved independently in various embodiments or may be combined in yetother embodiments further details of which can be seen with reference tothe following description and drawings.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1A is an illustration of an example aircraft that may embody theexamples described herein.

FIG. 1B illustrates an example aircraft engine having an example bleedair system disclosed herein.

FIG. 2 illustrates another aircraft engine having an example bleed airsystem disclosed herein.

FIG. 3 illustrates another aircraft engine having another example bleedair system disclosed herein.

FIG. 4 illustrates another aircraft engine having another example bleedair system disclosed herein.

FIG. 5 illustrates another aircraft engine having another example bleedair system disclosed herein.

FIG. 6 illustrates another example aircraft engine disclosed herein.

FIG. 7 illustrates another aircraft engine having another example bleedair system disclosed herein.

FIG. 8 is a flowchart representative of an example method that may beperformed by the example bleed air systems of FIGS. 1B and 2-7.

FIG. 9 is a flowchart illustrating a method of implementing the examplebleed air systems of FIGS. 1B and 2-7.

FIG. 10 illustrates another aircraft engine having another example bleedair system disclosed herein.

FIG. 11 is a flowchart illustrating a method of implementing the examplebleed air systems of FIG. 10.

FIG. 12 is a flowchart illustrating a method of implementing the examplebleed air system of FIG. 10.

Wherever possible, the same reference numbers will be used throughoutthe drawing(s) and accompanying written description to refer to the sameor like parts. As used in this patent, stating that any part (e.g., alayer, film, area, or plate) is in any way positioned on (e.g.,positioned on, located on, disposed on, or formed on, etc.) anotherpart, means that the referenced part is either in contact with the otherpart, or that the referenced part is above the other part with one ormore intermediate part(s) located therebetween. Stating that any part isin contact with another part means that there is no intermediate partbetween the two parts.

DESCRIPTION

Engine bleed air is typically provided by a compressor of an aircraftengine to power various systems of the aircraft. For example, bleed airis often used to power an environmental control system (ECS) and/or athermal anti-icing system of the aircraft. The bleed air is bled from acompressor of the aircraft engine via a bleed port in a housing of acompressor. However, bleed air pressures vary greatly with operatingconditions such as, for example, engine speed, operating altitude, etc.

To ensure the bleed air has sufficient pressure and/or temperature topower various systems of the aircraft, the bleed air is often extractedfrom a first bleed port (e.g., a low-pressure bleed port) of acompressor that provides sufficient pressure for the systems.Additionally, bleed air may also be provided via a high-pressure bleedport when the pressure of the low-pressure bleed air is insufficient tosupply the systems. For example, bleed air is often extracted from aneighth stage compressor of an aircraft engine during high engine speedsand from a fifteenth stage compressor during high altitude and/or lowengine speed operations. Thus, much of the energy spent by the engine toproduce the bleed air may be wasted if not completely used.

Additionally, the bleed air extracted from the engine often exceeds atemperature threshold of the aircraft systems utilizing the bleed air.Thus, the bleed air is cooled prior to supplying the bleed air to, forexample, the ECS. To reduce the bleed air temperature, commercialaircrafts typically employ a precooler (e.g., an air-to-air heatexchanger) through which bleed air passes and which is typically locatedon a pylon adjacent to the engine. A fan operated by the engine of theaircraft provides cool air to the precooler to cool the bleed air priorto supplying the bleed air to the systems of the aircraft. The fan airoften is dumped overboard after flowing through the precooler. Thus,cooling the bleed air via the fan often reduces the efficiency of theaircraft engine. Additionally, the precooler typically has a relativelylarge dimensional envelope, which adds extra weight and requires a fanair scoop and exhaust that produce drag. Thus, the relatively largedimensional envelope of the precooler can also affect the efficiency ofthe aircraft engine. Additionally or alternatively, an inlet port of theprecooler is positioned downstream and/or oriented opposite the enginebleed air port that provides the bleed air to the precooler. Thus, thebleed air is often piped to the inlet of the precooler using tight-bendelbows, which cause energy loss in the bleed air.

In some known examples, compressed air to the various systems of theaircraft is provided via electrically driven compressors. However,electrically driven compressors may not be efficient for relativelysmaller aircrafts. In other known examples, a bleed air system employs aturbo-compressor that receives ambient air from an atmospheric inlet.However, the atmospheric inlet produces drag. Additionally, theatmospheric inlet is often susceptible to icing and, thus, requires ananti-icing system that increases costs and system complexity. Further,the compressor may have to be relatively large to produce a pressurechange sufficient to power the systems of an aircraft.

Example bleed air systems and related methods disclosed herein employ aturbo-compressor to provide compressed or pressurized air to varioussystems of an aircraft such as, for example, an environmental controlsystem (ECS), a thermal anti-icing system (e.g., a wing and/or engine,anti-icing system), a pneumatic supply system (to supply pneumaticdevices), and/or any other system of the aircraft that requires use ofcompressed air. Unlike known systems, the example bleed air systemsdisclosed herein receive relatively lower pressure bleed air (e.g., froma fifth stage compressor) than known bleed air systems such as thosedescribed above. As a result, less energy is required from the engine toproduce the bleed air. Further, unlike known systems that employ aturbo-compressor, the example bleed air systems and related methodsdescribed herein enable use of a relatively smaller turbo-compressor.

Employing an example turbo-compressor system disclosed hereinsignificantly reduces an amount of high-pressure bleed air (or bleed airhaving relatively higher pressure) needed to satisfy the demand of anenvironmental control system of an aircraft. More specifically, thebleed air systems and related methods disclosed herein use bleed airhaving a relatively lower pressure and/or temperature to power systemsof an aircraft. For example, some example bleed air systems and relatedmethods disclosed herein employ a turbo-compressor that extracts bleedair from a low-pressure bleed port of a low-pressure compressor stage(e.g., a fifth stage). In other words, the example bleed air systems andrelated methods disclosed herein extract bleed air from a stage of acompressor that has a relatively lower pressure than the bleed airtypically extracted by known systems. By extracting bleed air from alower compressor stage of the aircraft engine, less energy in the bleedair is wasted, which significantly reduces the specific fuel consumptionof the engine.

Thus, with the examples disclosed herein, bleed air is extracted fromthe engine having relatively lower energy (e.g., temperature) than, forexample, an amount of energy in the bleed air extracted in knownsystems. More specifically, because the turbo-compressor employed by theexamples disclosed herein can boost the pressure of the bleed air, bleedair having a relatively lower pressure may be extracted from the engine,requiring less energy from the engine to produce sufficientlypressurized bleed air. In particular, the turbo-compressor increases thepressure of the low-pressure bleed air to a pressure sufficient for useby various systems of the aircraft. As a result, bleed air having lessenergy (e.g., a relatively lower pressure and/or temperature) may beextracted from the engine (e.g., a core of an engine) than knownsystems. Extracting bleed air having relatively less energy results inless wasted energy, thereby significantly increasing the fuel efficiencyof an aircraft engine (e.g., a turbofan engine).

In some examples, a turbo-compressor disclosed herein is capable ofreceiving bleed air at different pressures and/or temperatures fromdifferent stages of an aircraft engine. For example, when an aircraftengine is operating at a relatively high thrust (e.g., during take-offand/or cruising altitudes), the bleed air from a low-pressure compressorof the aircraft engine may have a higher pressure and/or temperaturethan, for example, the bleed air from the low-pressure compressor whenthe engine is operating at a relatively low thrust (e.g., when idle). Asa result, bleed air having different pressures and/or temperatures maybe utilized during different flight conditions to meet pressure and/ortemperature demands of an aircraft control system (e.g., anenvironmental control system ECS, an anti-icing system, etc.), therebyreducing a dimensional envelope (e.g., size and weight) of theturbo-compressor.

FIG. 1A illustrates an example commercial aircraft 100 having anaircraft engine 102 (e.g., turbofan engines) that may embody aspects ofthe teachings of this disclosure. FIG. 1B is a sectional view of theexample aircraft engine 102 of FIG. 1A. Each engine 102 of the aircraft100 may employ a dedicated bleed air system 104 and/or may employ acommon bleed air system 104. Further, the example bleed air system 104of FIG. 1B does not employ a precooler. The bleed air system 104 of FIG.1B provides compressed or pressurized air to an aircraft system 106 suchas, for example, an environmental control system 108 (ESC), a thermalanti-icing system 110 (e.g., an engine and/or wing anti-icing system),etc.

Turning in detail to FIG. 1B, the example bleed air system 104 employs aturbo-compressor 112 having a compressor 114 and a turbine 116. As shownin FIG. 1B, the turbo-compressor 112 is disposed within a nacelle 118 ofthe engine 102. Although the turbo-compressor 112 is disposed within thenacelle 118 as shown in FIG. 1B, in some examples, the turbo-compressor112 may be disposed at a remote location relative to the nacelle 118 orany other suitable location of the aircraft 100.

The compressor 114 of the turbo-compressor 112 receives low-pressurebleed air from a low-pressure source 120 (e.g., a low-pressure stage ofa compressor) of the engine 102. The compressor 114 compresses the bleedair and discharges the compressed bleed air to the ECS 108 of theaircraft 100. The turbine 116 receives high-pressure bleed air from ahigh-pressure source 122 (e.g., a high-pressure stage of a compressor)of the engine 102 providing bleed air having a higher pressure than thebleed air provided by the low-pressure source 120. Thus, the bleed airprovided to the turbine 116 has a relatively higher pressure and/ortemperature than the pressure and/or temperature of the bleed airprovided to the compressor 114. The high-pressure bleed air operates theturbine 116 which, in turn, operates the compressor 114. The turbine 116discharges the bleed air at a relatively lower pressure and/ortemperature, which is used to supply the thermal anti-icing system 110of the aircraft 100.

Further, the bleed air system 104 of the illustrated example includes atemperature boost or ejector 124 (e.g., an ejector conduit). The ejector124 boosts the temperature of the bleed air discharged from the turbine116 if the temperature of the bleed air is less than a minimumtemperature needed for thermal anti-icing operations. One or morecontrol systems, temperature sensors, valves, etc. may be employed tocontrol the flow of bleed air to the anti-icing system 110 via theejector 124.

FIG. 2 is a partial cutaway view of an engine 200 of an aircraft (e.g.,the aircraft 100) implemented with an example bleed air system 202 inaccordance with the teachings of this disclosure. Referring to FIG. 2,the engine 200 is a gas turbine having a fan 204 that draws air into anintake section 206 and a compressor 208. The compressor 208 may includemultiple compressor sections. For example, as shown, the compressor 208of the illustrated example is a dual-axial compressor that includes afirst compressor 210 and a second compressor 212. Additionally, each ofthe compressors 210 and 212 includes various different compressor stagesthat progressively increase the pressure of the air as the air flowsfrom the intake section 206 to a combustion chamber 214. For example,the first compressor 210 is a low-pressure compressor that providesrelatively low pressure air and the second compressor 212 is ahigh-pressure compressor that provides relatively high pressure air. Thecompressors 210 and 212 are coupled to respective shafts 216 and 218.The first shaft 216 is operatively coupled to a low-pressure turbine 220and the second shaft 218 is operatively coupled to a high-pressureturbine 222.

After exiting the high-pressure compressor 212, the highly pressurizedair is provided to the combustion chamber 214, where fuel is injectedand mixed with the highly pressure air and ignited. The high energyairflow exiting the combustion chamber 214 turns blades 224 of theturbines 220 and 222, which are coupled to respective ones of the shafts216 and 218. Rotation of the shafts 216 and 218 turns vanes 226 of thecompressors 210 and 212. The heated air is exhausted via a nozzle (notshown) where it mixes with cool air, provided by the fan 204, thatbypasses the engine core to produce forward thrust.

To supply various systems of an aircraft with pressurized air, theexample bleed air system 202 diverts or extracts some of the compressedair (e.g., bleed air) from the compressor 208 prior to the combustionchamber 214. In the illustrated example, the bleed air system 202includes a turbo-compressor 230 having a compressor 232 and a turbine234. In particular, the compressor 232 of the illustrated example has acompressor inlet 236 fluidly coupled to a bleed port 238 of thelow-pressure compressor 210. A conduit or passageway 240 fluidly couplesthe bleed port 238 of the low-pressure compressor 210 and the compressorinlet 236. The compressor 232 of the turbo-compressor 230 compresses thebleed air extracted from the low-pressure compressor 210 to provide arelatively higher pressure bleed air at a compressor outlet 242. Theexample bleed air system 202 does not employ a precooler. Thus, unlikeknown systems, the example bleed air system 202 described herein doesnot dump or waste air provided by the fan 204, thereby increasing theefficiency of the engine 200.

To prevent back flow pressure to the low-pressure compressor 210, thebleed air system 202 of the illustrated example employs a flow controlmember or back-flow preventer 244 (e.g., a check valve). The flowcontrol member 244 is disposed between the bleed port 238 of thelow-pressure compressor 210 and the compressor inlet 236 to preventpressurized airflow toward the low-pressure compressor 210 of the engine200 during a pressure surge produced by the turbo-compressor 230. Inother words, if the compressor 232 generates a back pressure at thecompressor inlet 236 that is greater than the pressure of the bleed airextracted from the low-pressure compressor 210 of the engine 200, theflow control member 244 prevents the higher pressure bleed air fromflowing from the compressor inlet 236 into the low-pressure compressor210.

In the illustrated example, the compressor inlet 236 is fluidly coupledto a low-pressure source such as, for example, a fifth stage of thelow-pressure compressor 210. Although in the illustrated example thecompressor inlet 236 receives bleed air from a dedicated compressorstage (e.g., a fifth stage) of the low-pressure compressor 210, thecompressor inlet 236 may alternatively receive bleed air from anotherstage of the low-pressure compressor 210 (e.g., a fourth stage, aneighth stage, etc.).

Additionally or alternatively, in some examples, the compressor inlet236 may be fluidly coupled to high-pressure source or a stage of thehigh-pressure compressor 212. In some such examples, the example bleedair system 202 may include dual-inlet bleed air ports. In that case, afirst port provides bleed air to the compressor inlet 236 from a firststage (e.g., a low-pressure stage such as the fifth stage) of thelow-pressure compressor 210 and a second port provides bleed air to thecompressor inlet 236 from a second stage (e.g., a low-pressure stagesuch as the eighth stage) of the low-pressure compressor 210 or a stageof the high-pressure compressor 212 (e.g., a high-pressure stage such asa fifteenth stage). In some such examples, one or more flow controlapparatus or valves may be employed to direct the bleed air from eitherone of the dual ports to the compressor inlet 236. For example, when theengine is operating at a relatively high thrust (e.g., during take-offand/or cruising altitudes), the flow control apparatus may direct bleedair from the low-pressure compressor 210. Alternatively, when the engineis operating at a relatively low thrust (e.g., when idle), the flowcontrol apparatus directs bleed flow to the compressor inlet 236 from ahigher stage of the low-pressure compressor 210 or the high-pressurecompressor 212.

As shown in FIG. 2, the compressor outlet 242 is fluidly coupled to acontrol system 246 of an aircraft. For example, a passageway or conduit248 fluidly couples the compressor outlet 242 to an environmentalcontrol system 250 (ECS). The ECS 250 conditions the bleed air providedby the compressor outlet 242 to a cabin pressure and/or cabintemperature. For example, the ECS 250 may include one or more ECS packs(e.g., an air cycle refrigeration system) that receive the bleed airfrom the turbo-compressor 230 and condition or regulate the bleed air tocabin pressure and/or temperature. A flow control member 252 (e.g., apressure reducing regulator) is disposed between the compressor outlet242 and the ECS 250. For example, the flow control member 252 may beconfigured to regulate the pressure of the bleed air at the compressoroutlet 242 to a pre-set or predetermined pressure value prior tosupplying the bleed air to the ECS 250. Additionally, the flow controlmember 252 may provide fluid flow shut-off (e.g., a pressure reducingshut-off valve or PRSOV).

The compressor outlet 242 of the illustrated example is also fluidlycoupled to a thermal anti-icing system 254. In the illustrated example,the thermal anti-icing system 254 is a wing anti-icing system. Thepassageway 248 fluidly couples the compressor outlet 242 to a passagewayor conduit 256 of the wing anti-icing system 254. A flow control member258 may be disposed between the compressor outlet 242 and the winganti-icing system 254. As shown, the flow control member 258 and/or thepassageway 256 are fluidly coupled to the passageway 248 downstream fromthe flow control member 252. However, in other examples, the flowcontrol member 258 and/or the passageway 256 may be fluidly coupled tothe passageway 248 upstream from the flow control member 252. The flowcontrol member 258 of the illustrated example is a pressure regulatorthat regulates the pressure of the bleed air to a pre-set pressuresuitable for the wing anti-icing system 254. Additionally, the flowcontrol member 258 may also be a shut-off valve (e.g., a PRSOV).

To drive the compressor 232, the turbo-compressor 230 employs theturbine 234. The turbine 234 is operatively coupled to the compressor232 via a shaft 260. To drive the turbine 234, a turbine inlet 262 isfluidly coupled to the high-pressure compressor 212 of the engine 200. Aconduit or passageway 264 fluidly couples a bleed port 266 of thehigh-pressure compressor 212 of the engine 200 and the turbine inlet262. For example, the bleed port 266 of the illustrated example fluidlycouples a fifteenth stage of the high-pressure compressor 212 to theturbine inlet 262. However, in other examples, the turbine inlet 262 maybe fluidly coupled to any other stage of the high-pressure compressor212. To regulate the pressure of the bleed air extracted from thehigh-pressure compressor 212 to a pre-set or predetermined pressurevalue prior to flowing into the turbine inlet 262, the bleed air system202 of the illustrated example employs a flow control member or pressureregulator 268. As shown in FIG. 2, the flow control member 268 isdisposed within the passageway 264 between the high-pressure compressor212 and the turbine inlet 262. Additionally, the flow control member 268may also be a shut-off valve (e.g., a PRSOV). High-pressure bleed airprovided by the high-pressure compressor 212 drives the turbine 234which, in turn, drives the compressor 232 via the shaft 260. The turbine234 converts thermal energy provided by the bleed air of thehigh-pressure compressor 212 into mechanical energy by expanding thehot, high-pressure bleed air provided by the high-pressure compressor212 to a lower temperature and pressure.

A turbine outlet 270 is fluidly coupled to a thermal anti-icing system272 of the aircraft 100. In the illustrated example of FIG. 2, thethermal anti-icing system 272 is an engine anti-icing system. A conduitor passageway 274 fluidly couples the turbine outlet 270 and the engineanti-icing system 272 to enable the bleed air at the turbine outlet 270to supply the engine anti-icing system 272. The engine anti-icing system272 of the illustrated example is configured to dump the bleed airexiting the turbine outlet 270. As shown in FIG. 2, a flow controlmember or pressure regulator 276 is disposed within the passageway 274between the turbine outlet 270 and the engine anti-icing system 272. Forexample, the flow control member 276 may regulate a pressure of thebleed air at the turbine outlet 270 prior to supplying the engineanti-icing system 272. Additionally, the flow control member 276 mayalso be a shut-off valve (e.g., a PRSOV). Although not shown, in otherexamples, the bleed air provided at the turbine outlet 270 may also beused to supply the wing anti-icing system 254.

In operation, the turbine 234 receives high-pressure bleed air from thehigh-pressure compressor 212 of the engine 200. The turbine 234 operatesthe compressor 232 and the bleed air exits the turbine outlet 270 tosupply the engine anti-icing system 272. The compressor 232 receiveslow-pressure bleed air from the low-pressure compressor 210 of theengine 200 and compresses the low-pressure bleed air to a relativelyhigher pressure bleed air to supply the ECS 250 and/or the winganti-icing system 254.

The bleed air system 202 may employ a control system 280 to operate thevarious flow control members 252, 258, 268 and 276. The control system280 of the illustrated example includes a microprocessor 282, aninput/output module 284, a comparator 286, and a valve controller 288.For example, a sensor 290 downstream of the flow control member 252 mayprovide a signal corresponding to a pressure of the bleed air prior toflowing to the ECS 250 to the microprocessor 282 via the input/outputmodule 284. The control system 280 may determine if the downstreampressure measured by the sensor 290 is within a pre-determined pressurerange for a given altitude and/or a given speed of the aircraft. Forexample, the comparator 286 may compare the signal provided by thesensor 290 with a pressure threshold or range provided, for example, bya look-up table. If the pressure is outside of the range, the controlsystem 280 may cause the flow control member 252 to adjust the bleed airpressure to within the predetermined pressure range. The control system280 may operate the other flow control members 258, 268 and 276 in asimilar manner. In some examples, the control system may cause the flowcontrol members 252, 258, 268 and 276 to operate between a closedposition to restrict fluid flow through the respective passageways 248,256, 264 or systems 250, 254, 272. Further, the control system 280 maybe part of a main control system 292 of the aircraft.

FIG. 3 is illustrates another example bleed air system 302 describedherein. Those components of the example bleed air system 302 that aresubstantially similar or identical to the components of the examplebleed air system 202 described above and that have functionssubstantially similar or identical to the functions of those componentswill not be described in detail again below. Instead, the interestedreader is referred to the above corresponding descriptions. Tofacilitate this process, similar reference numbers will be used for likestructures.

Referring to FIG. 3, the example bleed air system 302 employs asupplemental bleed air bypass system 304. In particular, thesupplemental bleed air bypass system 304 fluidly couples bleed air fromthe high-pressure compressor 212 to the compressor outlet 242. As shownin FIG. 3, the supplemental bleed air bypass system 304 includes aconduit or passageway 306 to fluidly couple a bleed port 308 of thehigh-pressure compressor 212 to the passageway 248 downstream from thecompressor outlet 242. For example, if a bleed air system (e.g., thebleed air system 202) of another engine 200 cannot provide adequatepressurized air to the ECS 250 and/or the wing anti-icing system 254,the supplemental bleed air bypass system 304 boosts the pressure and/ortemperature to the ECS 250 and/or the wing anti-icing system 254. Inthis manner, the turbo-compressor 230 can be sized smaller than aturbo-compressor that may otherwise be needed to support two or moreaircraft systems (ECS and/or anti-icing systems). Thus, the supplementalbleed air bypass system 304 enables the turbo-compressor 230 to have arelatively smaller dimensional envelope.

The supplemental bleed air bypass system 304 may employ a flow controlapparatus 310 such as, for example, a digital pressure regulatedshut-off valve. The flow control apparatus 310 of the illustratedexample is disposed in the passageway 306 between the bleed port 308 ofthe high-pressure compressor 212 and the passageway 248 downstream fromthe compressor outlet 242. A sensor 312 may be employed to detect ormeasure the pressure and/or temperature of the bleed air downstream fromthe compressor outlet 242 and generate a signal corresponding to themeasured pressure and/or temperature to the control system 280. Thecontrol system 280 receives the signal from the sensor 312 and may beconfigured to compare the measured pressure and/or temperature to athreshold value via the comparator 286. If the pressure and/ortemperature of the bleed air are less than a threshold value, then thecontrol system 280 may cause the flow control apparatus 310 to move toan open position to allow bleed air from the high-pressure compressor212 to flow to the passageway 248 downstream of the compressor outlet242. Alternatively, the control system 280 may operate the flow controlapparatus 310 (e.g., move it to an open position) if the control system280 receives a signal (e.g., from a main control system 292) indicatingthat a bleed air system of another engine of the aircraft is notproviding adequately pressurized bleed air.

FIG. 4 illustrates another example bleed air system 402 describedherein. Those components of the example bleed air system 402 that aresubstantially similar or identical to the components of the examplebleed air system 202 described above and that have functionssubstantially similar or identical to the functions of those componentswill not be described in detail again below. Instead, the interestedreader is referred to the above corresponding descriptions. Tofacilitate this process, similar reference numbers will be used for likestructures. Also, although not shown, the example bleed air system 402may also employ the supplemental bleed air bypass system 304 of FIG. 3.

The bleed air system 402 of FIG. 4 includes a mix-flow bypass system404. The mix-flow bypass system 404 of the illustrated example fluidlycouples the turbine outlet 270 and the compressor outlet 242 to enablethe bleed air from the turbine outlet 270 to mix with bleed air providedat the compressor outlet 242. In particular, the mix-flow bypass system404 enables the bleed air at the turbine outlet 270 to mix with thebleed air downstream from the compressor outlet 242 when the respectivepressures of the bleed air at the turbine outlet 270 and the compressoroutlet 242 are substantially equal. As a result, a lesser amount ofbleed air from the turbine outlet 270 is dumped overboard via the engineanti-icing system 272 during operation.

In the illustrated example of FIG. 4, a first passageway or conduit 406fluidly couples the turbine outlet 270 and the passageway 248 downstreamof the compressor outlet 242 and a second passageway or conduit 408fluidly couples the turbine outlet 270 and the engine anti-icing system272. A flow control apparatus or valve 410 such as, for example, athree-way pressure regulating valve is disposed between the turbineoutlet 270, the compressor outlet 242 and the engine anti-icing system272. In operation, the flow control member 410 receives bleed air fromthe turbine outlet 270 via a passageway 412 and the flow control member410 directs the bleed air to the passageway 406 and/or the passageway408.

A first sensor 414 such as, for example, a pressure sensor detects ormeasures the pressure (e.g., the static pressure) of the bleed air atthe compressor outlet 242 and a second sensor 416 such as, for example,a pressure sensor detects or measures the pressure (e.g., the staticpressure) of the bleed air at the turbine outlet 270. The sensors 414and 416 provide respective signals to the control system 280 of theaircraft 100 via the input/output module 284. The control system 280,via the comparator 286, compares the pressure values corresponding tothe signal values provided by the sensors 414 and 416. Themicroprocessor 282 commands the valve controller 288 to operate the flowcontrol apparatus 410 based on the compared pressure values. Forexample, the flow control apparatus 410 may be a digital pressureregulated shut-off valve.

In operation, if a comparison between the measured pressure of the bleedair at the compressor outlet 242 and the measured pressure of the bleedair at the turbine outlet 270 indicate that the pressures aresubstantially equal, the control system 280 sends a signal to the valvecontroller 288, which operates the flow control apparatus 410 to allowbleed air at the turbine outlet 270 to flow to the compressor outlet242. In this manner, the bleed air at the turbine outlet 270 mixes withthe bleed air at the compressor outlet 242, resulting in less bleed airbeing dumped overboard during operation. Additionally or alternatively,the flow control apparatus 410 may be configured to regulate or reducethe pressure of the bleed air at the turbine outlet 270 to substantiallymatch the pressure of the bleed air at the compressor outlet 242 priorto mixing. Further, during a mixing operation, the flow controlapparatus 410 may be configured to also provide bleed air from theturbine outlet 270 to engine anti-icing system 272 via the passageway408. Thus, the mix-flow bypass system 404 may direct all of the bleedair at the turbine outlet 270 to the compressor outlet 242 or only aportion of the bleed air at the turbine outlet 270 to the compressoroutlet 242.

Unbalanced pressures between the turbine outlet 270 and the compressoroutlet 242 may alter or affect the operation of the compressor 232and/or the turbine 234. For example, a pressure (e.g., a back pressure)provided by the bleed air at the compressor outlet 242 that is greaterthan the pressure of the bleed air at the turbine outlet 270 may affectthe operation of the turbine 234, and a pressure (e.g., a back pressure)provided by the bleed air at the turbine outlet 270 that is greater thanthe pressure of the bleed air at the compressor outlet 242 may affectthe operation of the compressor 232. Thus, if the sensors 414 and 416provide signals indicating that the respective pressures of the bleedair at the compressor outlet 242 and the turbine outlet 270 aredifferent or outside a predetermined pressure range (e.g., are lessthan, greater than or outside of a certain pressure range), the controlsystem 280 causes the flow control apparatus 410 to prevent the bleedair at the turbine outlet 270 from flowing to the compressor outlet 242via the passageway 406 and the control system 280 causes the flowcontrol apparatus 410 to allow the bleed air at the turbine outlet 270to flow to the engine anti-icing system 272.

As noted above, the example system of FIG. 4 reduces an amount of bleedair that is dumped overboard during operation. As a result, theefficiency of the engine 200 increases because the amount of energystored in the bleed air is used in various systems (e.g., the ECS 250,the wing anti-icing system 254, etc.) of the aircraft and not dumpedoverboard, wasted or lost.

FIG. 5 illustrates another example bleed air system 502 describedherein. Those components of the example bleed air system 502 that aresubstantially similar or identical to the components of the examplebleed air systems 202 and 402 described above and that have functionssubstantially similar or identical to the functions of those componentswill not be described in detail again below. Instead, the interestedreader is referred to the above corresponding descriptions. Tofacilitate this process, similar reference numbers will be used for likestructures.

The example bleed air system 502 of FIG. 5 includes an anti-icing boostbypass system 504. The anti-icing boost bypass system 504 of theillustrated example provides a higher temperature and/or higher pressurefluid to supplement or supply a thermal anti-icing system such as, forexample, the engine anti-icing system 272. For example, during certainoperating conditions of the engine 200 (e.g., when the engine 200 isidle), the engine 200 may operate with relatively lower power settings.As a result, a pressure and/or temperature of the bleed air at theturbine outlet 270 may be lower than a minimum temperature needed forthe engine anti-icing system 272. To increase the pressure and/or thetemperature of the bleed air at the turbine outlet 270, the anti-icingboost bypass system 504 boosts the temperature of the bleed air providedto the engine anti-icing system 272.

To boost the temperature of the bleed air supplied to the engineanti-icing system 272, the anti-icing boost bypass system 504 allowsbleed air from the high-pressure bleed port 266 to flow directly to theengine anti-icing system 272 via a passageway or conduit 506. To controlthe flow of fluid through the passageway 506, the anti-icing boostbypass system 504 employs a flow control apparatus or regulator valve508 disposed between the bleed port 266 and the passageway 506. Forexample, the flow control apparatus 508 is a digital pressure regulatingvalve. The passageway 264 fluidly couples the flow control apparatus 508to the high-pressure compressor 212 and the passageway 506 fluidlycouples the flow control apparatus 508 to the passageway 408 of theengine anti-icing system 272. As a result, the anti-icing boost bypasssystem 504 provides a higher temperature bleed air to the engineanti-icing system 272 without impacting or affecting the temperature ofthe bleed air used for a passenger cabin of the aircraft and provided bythe ECS 250. The anti-icing boost bypass system 504 may be an ejectorconduit.

Additionally or alternatively, the anti-icing boost bypass system 504may be fluidly coupled to the ECS 250 and/or the wing anti-icing system254 to provide higher pressure and/or temperature bleed air from thehigh-pressure compressor 212 to the compressor outlet 242 via the flowcontrol apparatus 410 and the passageways 406, 408 and 506. Such anexample may include one or more pressure regulators and/or flow controlapparatus to regulate the pressure of the bleed air from the anti-icingboost bypass system 504 to a pressure of the bleed air downstream of thecompressor outlet 242. Alternatively, although not shown, the examplebleed air system 502 of FIG. 5 may employ the supplemental bleed airbypass system 304 of FIG. 3 instead of the anti-icing boost bypasssystem 504 providing the bleed air to the ECS 250 and/or the winganti-icing system 254.

To determine whether a pressure and/or temperature boost to the engineanti-icing system 272, the ECS 250 and/or the wing anti-icing system 254is needed, the anti-icing boost bypass system 504 employs the controlsystem 280. The control system 280 may employ one or more sensors 510,512, 514 such as, for example, temperature sensor (e.g., a thermocouple)or a pressure sensor to measure the pressure and/or the temperature atthe turbine outlet 270 flowing to the engine anti-icing system 272and/or at the compressor outlet 242. The pressure/temperature sensors510, 512, 514 provide a signal to the control system 280 representativeof the pressure and/or the temperature of the bleed air in thepassageways 248, 406 and/or 408. For example, if the control system 280determines that the pressure and/or the temperature of the bleed air inthe passageway 408 is below a minimum threshold pressure and/ortemperature value, then the control system 280 causes the valvecontroller 288 to operate the flow control apparatus 508 to allow thebleed air from the high-pressure compressor 212 to flow to the engineanti-icing system 272 via the passageways 506 and 408. If the pressureand/or the temperature of the bleed air in the passageway 408 aregreater than the temperature threshold value, the control system 280causes the flow control apparatus 508 to prevent the bleed air from thehigh-pressure compressor 212 from flowing to the engine anti-icingsystem 272 via the passageway 506. Further, if the control system 280determines that a pressure and/or temperature boost is needed downstreamof the compressor outlet 242, then the control system 280 may beconfigured to cause the flow control apparatus 410 to allow fluid flowfrom the passageway 506 and 406.

Alternatively, in the example bleed air systems 202, 302, 402, 502described above (and 602 and 702 described below), the control system280 may operate the various flow control members 252, 258, 268, 276,310, 410 and/or 508 based on the operating conditions of the engine 200.For example, during operation, different engine speeds and/or differentaltitudes significantly vary the pressure and/or temperature of thebleed air provided by the low-pressure compressor 210 and thehigh-pressure compressor 212. For example, during high-thrust operation,the pressure of the bleed air provided both the low-pressure compressor210 and the high-pressure compressor 212 is significantly higher thanthe pressure of the bleed air provided when the engine 200 is at idle.Therefore, the control system 280 may receive indication of theoperating parameters or conditions (from the various sensors 290, 312,414, 416, 510, 512, 514, 616 and/or main control system 292) of theengine 200 and may operate the various flow control members 252, 258,268, 310, 410 and/or 508 accordingly. For example, if the control system280 determines that the engine 200 is idle, the control system 280 maycause the flow control apparatus 508 to move to an open position toallow higher temperature and/or higher pressure bleed air to flow fromthe high-pressure compressor 212 to the engine anti-icing system 272.

Also, because the anti-icing boost bypass system 504 employs the flowcontrol apparatus 508, the engine anti-icing flow control apparatus 276(e.g., a pressure regulator) may be removed. Otherwise, as shown in FIG.5, the engine anti-icing flow control apparatus 276 provides a redundantor safety valve apparatus. In yet other examples, the mix-flow bypasssystem 404 described in FIG. 4 may be removed from the bleed air system502 described in FIG. 5.

FIG. 6 illustrates another example bleed air system 602 describedherein. Those components of the example bleed air system 602 that aresubstantially similar or identical to the components of the examplebleed air system 202 described above and that have functionssubstantially similar or identical to the functions of those componentswill not be described in detail again below. Instead, the interestedreader is referred to the above corresponding descriptions. Tofacilitate this process, similar reference numbers will be used for likestructures. Also, although not shown, the example bleed air system 602may employ the supplemental bleed air bypass system 304 of FIG. 3.Further, the intercooler 604 of the FIG. 6 may be employed with theother example bleed air systems 104, 202, 302, 402, 502, and 702disclosed herein.

Unlike the bleed air systems 104, 202, 302, 402, 502 and 702 disclosedherein, the example bleed air system 602 employs an intercooler 604(e.g., an air-to-air heat exchanger). The bleed air system 602 of FIG. 6includes the turbo-compressor 230 having the compressor 232 and theturbine 234. In particular, the compressor inlet 236 of the compressor232 is fluidly coupled to the bleed port 238 of the low-pressurecompressor 210. Prior to the compressor inlet 236 receiving the bleedair from the bleed port 238, the bleed air flows through the intercooler604 between an intercooler inlet 606 and an intercooler outlet 608. Tocool the bleed air, the intercooler 604 receives air provided by the fan204 of the engine 200. The fan air flows through the intercooler 604between an inlet 610 and an outlet 612. The cooler fan air flowingbetween the inlet 610 and the outlet 612 reduces the temperature of thebleed air flowing between the intercooler inlet 606 and the intercooleroutlet 608.

To control the temperature of the bleed air at the intercooler outlet608, the bleed air system 602 employs a fan air valve 614. The fan airvalve 614 varies the cooling fan airflow provided by the fan 204, andthereby controls the air temperature of the bleed air at the pre-cooleroutlet 608. Alternatively, the intercooler 604 may receive ambient airfrom an atmospheric inlet.

The bleed air system 604 of the illustrated example employs a sensor 616(e.g., a temperature sensor) downstream from intercooler outlet 608 tosense or detect the temperature of the bleed air. The sensor 616 sendsthe signal corresponding to the temperature of the bleed air at theintercooler outlet 608 to the control system 280. The control system 280determines if the temperature value at the intercooler outlet 608 is ator within a threshold value or range. For example, if the control system280 determines (e.g., via the comparator) that the temperature of thebleed air is greater than a threshold temperature value, then thecontrol system 280 instructs the valve controller 288 to command the fanair valve 614 to allow more fan air flow into the inlet 610. If thetemperature of the bleed air at the intercooler outlet 608 is less thanthe threshold value, then the control system 280 commands the valvecontroller 288 to reduce the amount of fan air flowing to the inlet 610.The fan air exiting the outlet 612 is dumped overboard. In someexamples, the fan air exiting the outlet 612 of the intercooler 604 maybe routed for use in other systems of the aircraft.

Unlike known pre-coolers, the intercooler 604 of the illustrated exampleemploys a substantially smaller dimensional envelope and weight becausethe intercooler 604 does not need to be sized to hold a minimum volumeas required by known bleed air systems employing a known pre-cooler.Further, the intercooler 604 enables provision of a relatively smallersized turbo-compressor (e.g., smaller than the turbo-compressor 230).Additionally or alternatively, the intercooler 604 significantly reducesa temperature of the bleed air at the compressor outlet 242 (e.g., adischarge temperature of the turbo-compressor compressor 230). As aresult, the intercooler 604 reduces the amount of cooling that istypically provided by the ECS 250. In this manner, the size and or thenumber of ECS pack heat exchangers of the ECS 250 can be reduced,resulting in less aircraft drag. By significantly reducing the bleed airtemperature at the compressor outlet 242, the intercooler 604 performssome of the cooling function that is typically provided by the ECS 250.

FIG. 7 is a partial cut-away view of another example aircraft engine 700having another example bleed air system 702 described herein. Thosecomponents of the example engine 700 and the bleed air system 702 thatare substantially similar or identical to the components of the engine200 and the example bleed air system 202 described above and that havefunctions substantially similar or identical to the functions of thosecomponents will not be described in detail again below. Instead, theinterested reader is referred to the above corresponding descriptions.To facilitate this process, similar reference numbers will be used forlike structures.

In the illustrated example, the bleed air system 702 employs theturbo-compressor 230 disposed within a housing 704 of the engine 700.The compressor 232 of FIG. 7 receives low-pressure bleed air from thelow-pressure bleed port 238 of the engine 700. More specifically, asshown in FIG. 7, the compressor inlet 236 receives bleed air from afifth stage of the low-pressure compressor 210 and provides compressedor pressurized air to, for example, the thermal anti-icing systems 254,272, an environmental control system 250 and/or any other system of theaircraft that operates via compressed air. In some examples, the bleedair system 702 may employ a dual inlet source to enable the compressor232 to receive bleed air from a high-pressure source (e.g., thehigh-pressure compressor 212) during a first period of operation (e.g.,when the engine 700 is idle) and enable the compressor 232 to draw bleedair from a low-pressure source during a second period of operation(e.g., during cruising altitudes).

In the illustrated example, the turbine 234, via the turbine inlet 262,receives high-pressure bleed air from the high-pressure turbine 222 ofthe engine 700. Although not shown, the turbine 234 may alternativelyreceive bleed air from a high-pressure compressor 212 (e.g., any one ofthe different stages). In some examples, the bleed air system 702 mayemploy a dual inlet source to enable the turbine 234 to receive bleedair from a high-pressure source (e.g., the high-pressure turbine 222)during a first period of operation (e.g., when the engine 700 is idle)and enable the turbine 234 to draw bleed air from a lower-pressure stageof the turbine 222 and/or the high-pressure compressor 212 during asecond period of operation (e.g., during take-off). As noted above, theturbine 234 receives the bleed air to operate the compressor 232 via theshaft 260.

In the illustrated example, the lower temperature bleed air dischargedfrom the turbine 234 is used for cooling. In general, the turbine outlet270 may be fluidly coupled to a system (e.g., a second or other systemof the aircraft 100) requiring cooling. For example, as shown in FIG. 7,the turbine outlet 270 of the illustrated example discharges into alow-pressure (LP) turbine casing 706. In particular, the cooled bleedair discharged from the turbine 234 can be used to cool turbine blades708, the casing 706, etc. Cooling the casing 706 and/or the blades 708,for example, maintains a proper clearance or spacing between the turbineblades 708 and the casing 706 (e.g., prevents shrinkage and/or expansiondue to elevated temperatures). In other examples, the turbine outlet 270may be fluidly coupled to cool other portions of the turbine 222, 224and/or any other component of the aircraft engine 700 or the aircraft.In some examples, the bleed air system 702 may be combined with anyoneof the example bleed air systems 104, 202, 302, 402, 502 and 602disclosed above. For example, the turbine outlet 270 may discharge bleedair for cooling and/or for use in the anti-icing system 272.

FIG. 8 is a flowchart representative of an example method that may beimplemented with an example bleed air system such as, for example, thebleed air systems 104, 202, 302, 402, 502, 602 and/or 702 of FIGS. 1Band 2-7 and/or a control system such as the control system 280 of FIGS.2-7. In this example, the method comprises a program for execution by aprocessor such as the processor 282 shown in the example control system280. The program may be embodied in software stored on a tangiblecomputer readable medium such as a CD-ROM, a floppy disk, a hard drive,a digital versatile disk (DVD), or a memory associated with theprocessor and/or the control system 280, but the entire program and/orparts thereof could alternatively be executed by a device other than theprocessor 282 and/or embodied in firmware or dedicated hardware.Further, although the example program is described with reference to theflowchart illustrated in FIG. 8, many other methods of implementing theexample control system 280 and/or the valve controller 288 mayalternatively be used. For example, the order of execution of the blocksmay be changed, and/or some of the blocks described may be changed,eliminated, or combined.

As mentioned above, the example method or process of FIG. 8 may beimplemented using coded instructions (e.g., computer readableinstructions) stored on a tangible computer readable medium such as ahard disk drive, a flash memory, a read-only memory (ROM), a compactdisk (CD), a digital versatile disk (DVD), a cache, a random-accessmemory (RAM) and/or any other storage media in which information isstored for any duration (e.g., for extended time periods, permanently,brief instances, for temporarily buffering, and/or for caching of theinformation). As used herein, the term tangible computer readable mediumis expressly defined to include any type of computer readable storageand to exclude propagating signals. Additionally or alternatively, theexample process of FIG. 8 may be implemented using coded instructions(e.g., computer readable instructions) stored on a non-transitorycomputer readable medium such as a hard disk drive, a flash memory, aread-only memory, a compact disk, a digital versatile disk, a cache, arandom-access memory and/or any other storage media in which informationis stored for any duration (e.g., for extended time periods,permanently, brief instances, for temporarily buffering, and/or forcaching of the information). As used herein, the term non-transitorycomputer readable medium is expressly defined to include any type ofcomputer readable medium and to exclude propagating signals.

For purposes of discussion, the example process 800 of FIG. 8 isdescribed in connection with the example control system 280 and theexample valve controller 288. In this manner, each of the exampleoperations of the example process 800 of FIG. 8 is an example manner ofimplementing a corresponding one or more operations performed by one ormore of the blocks of the example control system 280 or the valvecontroller 288 of FIGS. 2-7.

Turning in detail to FIG. 8, the control system 280 monitors a systemparameter of a bleed air system (block 802). For example, the controlsystem 280 may be configured to monitor one or more system parameters ofthe example bleed air systems 104, 202, 302, 402, 502, 602 and/or 702disclosed herein. To monitor a system parameter, the control system 280receives one or more signals from the sensors 290, 312, 414, 416, 510,512, 514, 616 and/or main control system 292. For example, a systemparameter may include a pressure of bleed air, a temperature of thebleed air, a speed of the engine 200, an altitude of the aircraft, etc.The control system 280 receives the signals via the input/output module284.

The control system 280 compares the system parameter to a thresholdvalue (block 804). For example, the control system 280 can compare, viathe comparator 286, the pressure and/or temperature of the bleed airreceived via the sensors 290, 312, 414, 416, 510, 512, 514, 616 torespective threshold values retrieved from a look-up table or storage orwith the other values received by the sensors 290, 312, 414, 416, 510,512, 514, 616. In some examples, the system parameter and/or thethreshold value may be based on other operating conditions of theaircraft.

The control system 280 may then determine if the system parametermeasured by, for example, the sensors 290, 312, 414, 416, 510, 512, 514,616 is within the threshold value (block 806). If the system parameteris within the threshold value, then the control system 280 returns toblock 802 to continue monitoring the system parameter. For example, ifthe pressure of the bleed air at the turbine outlet 270 is substantiallyequal to the pressure at the compressor outlet 242, the control system280 may cause the flow control member 410 to remain in an open positionto enable flow via the mix-flow bypass system 404.

If the control system 280 determines that the system parameter is notwithin the threshold value, then the control system 280 commands thevalve controller 288 to operate a flow control member (block 808). Forexample, if the pressure of the bleed air at the turbine outlet 270 isnot substantially equal to the pressure at the compressor outlet 242,the control system 280 may cause the flow control member 410 to move toa closed position to prevent flow via the mix-flow bypass system 404.

FIG. 9 is a flowchart of an example method of implementing the examplebleed air systems 202, 302, 402, 502, 602 and/or 702 in accordance withthe teachings of this disclosure. While an example manner ofimplementing the example bleed air systems, one or more of the blocksand/or processes illustrated in FIG. 8 may be combined, divided,re-arranged, omitted, eliminated and/or implemented in any other way.Further still, the example method of FIG. 9 may include one or moreprocesses and/or blocks in addition to, or instead of, those illustratedin FIG. 9, and/or may include more than one of any or all of theillustrated processes and/or blocks. Further, although the examplemethod 900 is described with reference to the flow chart illustrated inFIG. 9, many other methods of implementing a bleed air system mayalternatively be used.

Referring to FIG. 9, the turbo-compressor 230 is operatively coupled tothe engine 200 of an aircraft (block 902). The compressor inlet 236 ofthe turbo-compressor 230 is fluidly coupled to a low-pressure bleed airsource via, for example, the bleed port 238 and the passageway 240(block 904). The compressor outlet 242 of the turbo-compressor 230 isfluidly coupled, via the passageway 248, to a first system of theaircraft such as, for example, the ECS 250, wing anti-icing system 254,etc. (block 906). The turbine inlet 262 of the turbo-compressor 230 isfluidly coupled to the high-pressure compressor 212 of the engine 200via the passageway 264 (block 908), and the turbine outlet 270 of theturbo-compressor 230 is fluidly coupled to the second control system ofthe aircraft such as, for example, the engine anti-icing system 272(block 910). Additionally or alternatively, the turbine outlet 270 ofthe turbo-compressor 230 may be coupled to a system of the aircraft 100requiring cooling such as, for example a low-pressure (LP) turbinecasing 706. (block 910).

In some examples, the high-pressure bleed air source or bleed port 266may be directly or fluidly coupled to the second control system 272 viaa boost bypass system 504 (block 912).

In some examples, the turbine outlet 270 and the compressor outlet 242may be fluidly coupled to mix bleed air based on a comparison of apressure of the bleed air at the turbine outlet 270 and the bleed air atthe compressor outlet 242 (block 914). For example, the mix-flow bypasssystem 404 may be employed to fluidly couple the turbine outlet 270 andthe compressor outlet 242.

In some examples, the high-pressure bleed air source (e.g., thehigh-pressure compressor 212) may be fluidly coupled to the compressoroutlet 242 to bypass the turbo-compressor 230 and supplement bleed airbased on a volume flow rate of the bleed air at the compressor outlet242 (block 916). For example, the high-pressure bleed air source 212 maybe fluidly coupled to the compressor outlet 242 via the supplementalbleed air bypass system 304.

FIG. 10 illustrates another example bleed air system 1000 describedherein. Those components of the example bleed air system 1000 that aresubstantially similar or identical to the components of the exampleengine 200 or the example bleed air system 202 described above and thathave functions substantially similar or identical to the functions ofthose components will not be described in detail again below. Instead,the interested reader is referred to the above correspondingdescriptions. To facilitate this process, similar reference numbers willbe used for like structures.

Referring to FIG. 10, the example bleed air system 1000 employsdual-inlet bleed air system 1002. The dual-inlet bleed air system 1002includes a first dual-inlet bleed air system 1004 associated with thecompressor 232 of the turbo-compressor 230 and a second dual-inlet bleedair system 1006 associated with the turbine 234 of the turbo-compressor230. In this manner, the compressor 232 is capable of receiving bleedair at different pressures and/or temperatures from different stages ofthe low-pressure compressor 210 and/or the high-pressure compressor 212and the turbine 234 is capable of receiving bleed air at differentpressures and/or temperatures from different stages of the high-pressurecompressor 212 and/or the low-pressure compressor 210. For example, thefirst dual-inlet bleed air system 1004 can provide bleed air to thecompressor 232 from a fifth stage of the low-pressure compressor 210 oran eighth stage of the high-pressure compressor 210. Likewise, thesecond dual-inlet bleed air system 1006 can provide bleed air to theturbine 234 from a fifteenth stage of the high-pressure compressor 212or the eighth stage of the high-pressure compressor 212.

As a result, bleed air having different pressures and/or temperaturesmay be utilized during different flight conditions to meet pressureand/or temperature demands of the control system 246 (e.g., the ECS 250,the wing anti-icing system 254, the engine anti-icing system 272, etc.),while reducing a dimensional envelope (e.g., size and weight) of theturbo-compressor 230. For example, when the engine 200 is operating at arelatively high thrust (e.g., during take-off and/or cruisingaltitudes), the bleed air from the low-pressure compressor 210 may havea higher pressure and/or temperature than, for example, the bleed airfrom the low-pressure compressor 210 when the engine is operating at arelatively low thrust (e.g., when idle). Thus, to ensure that thepressure and/or temperature of the bleed air from the low-pressurecompressor 210 is sufficient to meet the demands of the control systems246, a turbo-compressor is typically sized (e.g., to provide a 10 to 1ratio) based on the pressure and/or temperature of the bleed air duringthe low thrust conditions. Thus, a turbo-compressor may have arelatively larger dimensional envelope to meet the demands of thecontrol system 246 during low thrust conditions than a turbo-compressorneeded to meet the demands of the control system 246 during high trustconditions. Thus, by using dual inlets, bleed air from a higher stage ofthe low-pressure compressor 210 and/or the high-pressure compressor 212may be provided to the turbo-compressor 230 during different flightconditions. By providing bleed air to the turbo-compressor 230 from ahigher pressure source of the compressors 210 and 212 during low thrustconditions (e.g., non-flight conditions or when the aircraft is idle),the turbo-compressor 230 (e.g., a 5 to 1 ratio) of the illustratedexample may be sized based on the higher pressure bleed air, resultingin a turbo-compressor 230 having a relatively smaller dimensionalenvelope, thereby reducing weight and costs associated with the aircraft100.

As shown in the illustrated example, the first dual-inlet bleed airsystem 1004 fluidly couples the compressor inlet 236 to the low-pressurecompressor 210 and the high-pressure compressor 212. As shown, the firstdual-inlet bleed air system 1004 includes a first inlet passageway 1008and a second inlet passageway 1010. For example, the first inletpassageway 1008 fluidly couples the compressor inlet 236 to thelow-pressure port 238 of the low-pressure compressor 210 and the secondinlet passageway 1010 fluidly couples the compressor inlet 236 to alow-pressure port or intermediate port 1012 of the high-pressurecompressor 212. For example, the low-pressure port 238 may provide fluidfrom a fifth stage of the low-pressure compressor 210 and theintermediate port 1012 may provide fluid from an eighth stage of thehigh-pressure compressor 212. As noted above, the pressure of the bleedair at the intermediate port 1012 is greater than the pressure of thebleed air at the low-pressure port 238.

Similarly, the second dual-inlet bleed air system 1002 fluidly couplesthe turbine 234 of the turbo-compressor 230 to different stages of thehigh-pressure compressor 212. As shown in the illustrated example, thesecond dual-inlet bleed air system 1006 includes a third inletpassageway 1014 and a fourth inlet passageway 1016. For example, thethird inlet passageway 1014 fluidly couples the turbine inlet 262 andthe high-pressure port 266 of the high-pressure compressor 212 and thefourth inlet passageway 1016 fluidly couples the turbine inlet 262 andthe intermediate port 1012 of the high-pressure compressor 212. Thehigh-pressure port 266 may provide bleed air from a fifteenth stage ofthe high-pressure compressor 212 and the intermediate port 1012 mayprovide bleed air from the eighth stage of the high-pressure compressor212. The pressure of the bleed air at the high-pressure port 266 isgreater than the pressure of the bleed air at the intermediate port1012.

As shown, the compressor inlet 236 and the turbine inlet 262 are fluidlycoupled to a common port (i.e., the intermediate port 1012). However, inother examples, the compressor inlet 236 or the turbine inlet 262 maynot be fluidly coupled to a common port. Thus, in some examples, thedual-inlet bleed air system 1002 provides bleed air to the compressor232 and the turbine 234 from a common stage of the low-pressurecompressor 210 and/or the high-pressure compressor 212 or from differentstages of the lower-pressure compressor 210 and/or the high-pressurecompressor 212.

To direct the bleed air from either one of the ports 238 and 1012 to thecompressor inlet 236 and/or either one of the ports 266 and 1012 to theturbine inlet 262, the dual-inlet bleed air system 1002 employs one ormore flow control members or valves 244, 268, 1018, 1020. For example,to control fluid flow between the low-pressure port 238 or theintermediate port 1012 and the compressor inlet 236, the firstdual-inlet bleed air system 1004 employs the flow control member 244(e.g., a check valve) coupled to the first inlet passageway 1008 and theflow control member 1018 (e.g., a pressure-relief shut-off valve)coupled to the second inlet passageway 1010. In particular, the flowcontrol member 244 is positioned between the low-pressure port 238 andthe compressor inlet 236 and the flow control member 1018 is positionedbetween the intermediate port 1012 and the compressor inlet 236. Inaddition, the first inlet passageway 1008 and the second inletpassageway 1010 converge into the passageway 240 downstream from therespective flow control members 244 and 1018 and upstream from thecompressor inlet 236.

Likewise, to control fluid flow between the turbine inlet 262 and thehigh-pressure port 266 or the intermediate port 1012, the seconddual-inlet bleed air system 1006 employs the flow control member 268coupled to the third inlet passageway 1014 and the flow control member1020 coupled to the fourth inlet passageway 1016. In particular, theflow control member 268 is positioned between the high-pressure port 266and the turbine inlet 262 and the flow control member 1020 is positionedbetween the intermediate port 1012 and the turbine inlet 262. The thirdinlet passageway 1014 and the fourth inlet passageway 1016 converge intothe passageway 264 downstream from the respective flow control members268 and 1020 and upstream from the turbine inlet 262.

The flow control member 244 of the first dual-inlet bleed air system1004 shown in the illustrated example is a check valve or back flowpreventer that allows fluid flow in one direction based on a pressuredifferential across the flow control member 244. More specifically, whenthe pressure of the bleed air upstream from the flow control member 244(e.g., at the port 238) is greater than the pressure of the bleed airdownstream from the flow control member 244 (e.g., at the compressorinlet 236), the flow control member 244 moves to an open position toallow bleed air to flow from the low-pressure port 238 to the compressorinlet 236. However, if the pressure downstream from the flow controlmember 244 (e.g., at the compressor inlet 236) is greater than thepressure of the bleed air upstream from the flow control member 244(i.e., at the low-pressure port 238), the flow control member moves to aclosed position to prevent fluid flow across the flow control member 244between the low-pressure port 238 and the compressor inlet 236.

Likewise, the flow control member 1020 of the second dual-inlet bleedair system 1006 shown in the illustrated example is a check valve orback flow preventer that allows fluid flow in one direction based on apressure differential across the flow control member 1020. Morespecifically, when the pressure of the bleed air upstream from the flowcontrol member 1020 (e.g., at the intermediate port 1012) is greaterthan the pressure of the bleed air downstream from the flow controlmember 1020 (e.g., at the turbine inlet 262), the flow control member1020 moves to a closed position preventing fluid flow across the flowcontrol member 1020 between the intermediate port 1012 and the turbineinlet 262. However, if the pressure of the bleed air upstream from theflow control member 1020 (e.g., at the turbine inlet 262) is less thanthe pressure of the bleed air downstream from the flow control member244 (i.e., at the intermediate port 1012), the flow control member 1020moves to an open position to allow bleed air to flow from theintermediate port 1012 to the turbine inlet 262.

The flow control members 244 and 1020 of the illustrated example movebetween the open and closed positions based on the operation of therespective flow control members 1018 and 268. The flow control members1018 and 268 of the respective first and second dual-inlet bleed airsystems 1004 and 1006 shown in the illustrated example are pressurerelief shut-off valves. Each of the flow control members 1018 and 268 iscontrollable or movable between an open position and a closed position.For example, when the flow control member 1018 of the first dual-inletbleed air system 1004 is an open position, bleed air from theintermediate port 1012 flows through the flow control member 1018 to thecompressor inlet 236 via the second inlet passageway 1010 and thepassageway 240. As a result, because the pressure of the bleed air atthe intermediate port 1012 is greater than the pressure of the bleed airat the low-pressure port 238, the pressure of the bleed air at thecompressor inlet 236 (e.g., at an outlet of the flow control member 244)is greater than the pressure of the bleed air at the low-pressure port238. Thus, the pressure differential across the flow control member 244causes the flow control member 244 to move to a closed position, therebypreventing fluid flow from the low-pressure port 238 to the compressorinlet 236 via the first inlet passageway 1008. However, when the flowcontrol member 1018 moves to a closed position, the bleed air from theintermediate port 1012 is prevented from flowing to the compressor inlet236. Thus, the pressure of the bleed air at the compressor inlet 236decreases. In turn, the pressure differential across the flow controlmember 244 causes the flow control member 244 to move to the openposition when the pressure downstream from the flow control member 244is less than the pressure of the bleed air upstream from the flowcontrol member 244, thereby allowing fluid flow from the low-pressureport 238 to the compressor inlet 236 via the first inlet passageway 1008and the passageway 240. In other words, when the flow control member1018 moves to an open position, the compressor inlet 236 receives bleedair from the intermediate port 1012 via the second inlet passageway 1010and when the flow control member 1018 moves to a closed position, thecompressor inlet 236 receives bleed air from the low-pressure port 238via the first inlet passageway 1008.

Similarly, when the flow control member 268 of the second dual-inletbleed air system 1006 moves to an open position, bleed air from thehigh-pressure port 266 flows through the flow control member 268 to theturbine inlet 262 via the third inlet passageway 1014 and the passageway264. As a result, the pressure of the bleed air upstream from the flowcontrol member 1020 (e.g., at the intermediate port 1012) is greaterthan the pressure of the bleed air downstream from the flow controlmember 1020 (e.g., at the intermediate port 1012). As a result, thepressure differential across the flow control member 1020 causes theflow control member 1020 to move to a closed position, therebypreventing fluid flow between the turbine inlet 262 and the intermediateport 1012 via the fourth inlet passageway 1016. When the flow controlmember 268 moves to a closed position, the bleed air from thehigh-pressure port 266 is prevented from flowing to the turbine inlet262. Thus, the pressure of the bleed air at the turbine inlet 262decreases. In turn, when a pressure downstream from the flow controlmember 1020 (e.g., at the turbine inlet 262) is less than the pressureof the bleed air upstream from the flow control member 1020 (e.g., atthe intermediate port 1012), the flow control member 1020 moves to anopen position allowing the bleed air from the intermediate port 1012 toflow to the turbine inlet 262 via the fourth inlet passageway 1016 andthe passageway 264. In other words, when the flow control member 268moves to a closed position, the turbine inlet 262 receives bleed airfrom the intermediate port 1012 and when the flow control member 268moves to an open position, the turbine inlet 262 receives bleed air fromthe high-pressure port 266.

The first dual-inlet bleed air system 1004 may be controlledindependently from the second dual-inlet bleed air system 1006. Forexample, the flow control members 1018 and 268 may be operatedindependently of each other such that the compressor inlet 236 canreceive bleed air from the intermediate port 1012 while the turbineinlet 262 receives bleed air from the high-pressure port 266; thecompressor inlet 236 may receive bleed air from the low-pressure port238 while the turbine inlet 262 receives bleed air from the intermediateport 1012; the compressor inlet 236 may receive bleed air from thelow-pressure port 238 while the turbine inlet 262 receives bleed airfrom the high-pressure port 266; or the compressor inlet 236 and theturbine inlet 262 may receive bleed air from the intermediate port 1012simultaneously. In other words, if one of the compressor 232 or theturbine 234 receives bleed air from the intermediate port 1012, theother one of the compressor 232 or the turbine 234 does not need toreceive the bleed air from the intermediate port 1012 even though thesecond inlet passageway 1010 is in fluid communication with the fourthinlet passageway 1016.

The control system 280 may operate the flow control members 1018 and268. As shown in the illustrated example, the bleed air system 1000employs the control system 280 to operate the dual-inlet bleed airsystem 1002. For example, to operate the dual-inlet bleed air system1002, the control system 280 may be configured to operate the flowcontrol members 1018 and 268 of the respective first and seconddual-inlet bleed air systems 1004 and 1006. In addition, the controlsystem 280 may be configured to operate the various flow control members252, 258 and 276 as indicated above.

The control system 280 of the illustrated example includes amicroprocessor 282, an input/output module 284, a comparator 286, and avalve controller 288. For example, a sensor 1022 may provide a signalcorresponding to a pressure of the bleed air to the microprocessor 282via the input/output module 284. The control system 280 may determine ifthe pressure value of the low-pressure port 238 measured by the sensor1022 is within a pre-determined pressure range and/or a pressure rangefor a given altitude and/or a given speed of the aircraft 100. Forexample, the comparator 286 may compare the signal provided by thesensor 1022 to a pressure threshold or range provided, for example, by alook-up table. Based on the comparison, the control system 280 maycommand the valve controller 288 to cause the flow control member 1018to move between an open position to allow bleed air to flow from theintermediate port 1012 to the compressor inlet 236 and a closed positionto allow bleed air to flow from the low-pressure port 238 to thecompressor inlet 236. Likewise, the control system 280 may determine,via for example a sensor 1024, if the pressure value of thehigh-pressure port 266 is greater than a predetermined pressure rangeand/or a pressure range for a given altitude and/or a given speed of theaircraft 100. For examples, the comparator 286 may compare a signalprovided by the sensor 1024 with a pressure threshold or range providedby, for example, the look-up table. Based on the comparison, the controlsystem 280 may command the valve controller 288 to cause the flowcontrol member 268 to move between a closed position to allow bleed airto flow from the intermediate port 1012 to the turbine inlet 262 and aclosed position to allow bleed air to flow from the high-pressure port266 to the turbine inlet 262. In some examples, the flow control members244 and/or 1020 may be pressure control valves (e.g., similar to theflow control members 1018 and 268) that are operated via the controlsystem 280.

In some examples, the dual-inlet bleed air system 1002 may be configuredto operate manually (e.g., based on the pressures of the bleed airduring high throttle or low throttle conditions). For example, the flowcontrol members 1018 and 268 may be pressure regulators (e.g., springoperated regulators such as spring to close valves) that move betweenopen and closed positions based on the pressure differential across theflow control members 1018 and 268 provided by the bleed air duringdifferent flight conditions (e.g., high thrust and low thrustconditions). For example, the flow control member 1018 may be configuredto move to a closed position when a pressure of the bleed air at theintermediate port 1012 is greater than a predetermined or pre-setpressure value (e.g., a pressure associated with a high thrustcondition) and move to an open position when the pressure of the bleedair at the intermediate port 1012 is less than the predetermined or thepre-set pressure value (e.g., a pressure indicative of a low thrustcondition). Also, the flow control member 268 may be configured to moveto a closed position when a pressure of the bleed air at thehigh-pressure port 266 is greater than a predetermined or pre-setpressure value (e.g., a pressure value associated with a high thrustcondition) and an open position when the pressure of the bleed air atthe high-pressure port 268 is less than the pre-determined or pre-setpressure value (e.g., a pressure value indicative of a low thrustcondition).

FIG. 11 is a flowchart representative of an example method 1100 that maybe implemented with an example dual-inlet bleed air system such as, forexample, the dual-inlet bleed air system 1002 of FIG. 10 and/or acontrol system such as the control system 280 of FIG. 10. In thisexample, the method comprises a program for execution by a processorsuch as the processor 282 shown in the example control system 280. Theprogram may be embodied in software stored on a tangible computerreadable medium such as a CD-ROM, a floppy disk, a hard drive, a digitalversatile disk (DVD), or a memory associated with the processor and/orthe control system 280, but the entire program and/or parts thereofcould alternatively be executed by a device other than the processor 282and/or embodied in firmware or dedicated hardware. Further, although theexample program is described with reference to the flowchart illustratedin FIG. 11, many other methods of implementing the example controlsystem 280 and/or the valve controller 288 may alternatively be used.For example, the order of execution of the blocks may be changed, and/orsome of the blocks described may be changed, eliminated, or combined.

As mentioned above, the example method or process of FIG. 10 may beimplemented using coded instructions (e.g., computer readableinstructions) stored on a tangible computer readable medium such as ahard disk drive, a flash memory, a read-only memory (ROM), a compactdisk (CD), a digital versatile disk (DVD), a cache, a random-accessmemory (RAM) and/or any other storage media in which information isstored for any duration (e.g., for extended time periods, permanently,brief instances, for temporarily buffering, and/or for caching of theinformation). As used herein, the term tangible computer readable mediumis expressly defined to include any type of computer readable storageand to exclude propagating signals. Additionally or alternatively, theexample process of FIG. 11 may be implemented using coded instructions(e.g., computer readable instructions) stored on a non-transitorycomputer readable medium such as a hard disk drive, a flash memory, aread-only memory, a compact disk, a digital versatile disk, a cache, arandom-access memory and/or any other storage media in which informationis stored for any duration (e.g., for extended time periods,permanently, brief instances, for temporarily buffering, and/or forcaching of the information). As used herein, the term non-transitorycomputer readable medium is expressly defined to include any type ofcomputer readable medium and to exclude propagating signals.

For purposes of discussion, the example process 1100 of FIG. 11 isdescribed in connection with the example dual-inlet bleed air system1002. In this manner, each of the example operations of the exampleprocess 1100 of FIG. 11 is an example manner of implementing acorresponding one or more operations performed by the example dual-inletbleed air system 1002 of FIG. 10.

Turning in detail to FIG. 11, the control system 280 monitors a systemparameter of the dual-inlet bleed air system 1002 (block 1102). Forexample, the control system 280 may be configured to monitor one or moresystem parameters of the example first dual-inlet bleed air system 1004or the second dual-inlet bleed air system 1006 disclosed herein. Tomonitor a system parameter, the control system 280 receives one or moresignals from the sensors 1022, 1024 and/or the main control system 292.For example, a system parameter may include a pressure and/or atemperature of the bleed air at the low-pressure port 238, thehigh-pressure port 266, a thrust condition of the engine 200, a speed ofthe aircraft 100 and/or an altitude of the aircraft 100 from the maincontrol system 292, etc. The control system 280 receives the signals viathe input/output module 284.

The control system 280 compares the system parameter to a thresholdvalue (block 1104). For example, the control system 280 can compare, viathe comparator 286, the measured pressure and/or temperature of thebleed air received via the sensors 1022 or 1024 to respective thresholdvalues retrieved from a look-up table or storage, or with the othervalues received by the sensors 1022 or 1024.

The control system 280 may then determine if the system parametermeasured by, for example, the sensors 1022 of the first dual-inlet bleedair system 1004 is within the threshold value (block 1106). If thesystem parameter is within the threshold value, then the control system280 causes the first dual-inlet bleed air system 1004 to provide bleedair from a first pressure source (block 1108). For example, the controlsystem 280 causes the flow control member 1018 to move to a closedposition to allow bleed air from the low-pressure port 238 to flow tothe compressor inlet 236.

If the system parameter is not within the threshold value, then thecontrol system 280 causes the first dual-inlet bleed air system 1004 toprovide bleed air from a second pressure source (block 1110). Forexample, the control system 280 causes the flow control member 1018 tomove to an open position to allow bleed air from the intermediate port1012 to flow to the compressor inlet 236.

The control system 280 may then determine if the system parametermeasured by, for example, the sensors 1024 of the second dual-inletbleed air system 1006 is within the threshold value (block 1112). If thesystem parameter is within the threshold value, then the control system280 causes the second dual-inlet bleed air system 1006 to provide bleedair from a third pressure source (block 1114). For example, the controlsystem 280 causes the flow control member 268 to move to an openposition to allow bleed air from the high-pressure port 266 to flow tothe turbine inlet 262.

If the system parameter is not within the threshold value, then thecontrol system 280 causes the second dual-inlet bleed air system 1006 toprovide bleed air from a fourth pressure source (block 1116). Forexample, the control system 280 causes the flow control member 268 tomove to a closed position to allow bleed air from the intermediate port1012 to flow to the turbine inlet 262.

In other examples, the system parameter may be based on a speed of theaircraft 100. For example, the control system 280 may receive a speedvalue representative of the aircraft speed. If the control systemdetermines that the speed value is greater than a threshold speed value(e.g., a speed representing in-flight or relatively high thrustconditions), then the control system 280 causes the flow control members1018 and 268 to move to their respective closed positions. If thecontrol system 280 determines that the speed value is less than thethreshold speed value, then the control system 280 causes the first andsecond flow control members 1018 and 268 to move to their respectiveopen positions.

In yet other examples, the control system 280 may operate the first andsecond flow control members 1018 and 268 based on system parameters(e.g., pressures and/or temperatures of the bleed air) upstream from thecompressor outlet 242 and/or the turbine outlet 270. For example, thecontrol system 280 may cause the flow control members 1018 and 268 tomove between their respective open and closed positions based on acomparison of the pressure and/or temperature of the bleed air at thecompressor outlet 242 or the turbine outlet 270 relative to apredetermined threshold pressure or temperature value.

FIG. 12 is a flowchart of an example method 1200 of implementing theexample bleed air system 1002 in accordance with the teachings of thisdisclosure. While an example manner of implementing the example bleedair systems, one or more of the blocks and/or processes illustrated inFIG. 12 may be combined, divided, re-arranged, omitted, eliminatedand/or implemented in any other way. Further still, the example methodof FIG. 12 may include one or more processes and/or blocks in additionto, or instead of, those illustrated in FIG. 12, and/or may include morethan one of any or all of the illustrated processes and/or blocks.Further, although the example method 1200 is described with reference tothe flow chart illustrated in FIG. 12, many other methods ofimplementing the bleed air system 1002 may alternatively be used.

Referring to FIG. 12, the turbo-compressor 230 is operatively coupled tothe engine 200 of the aircraft 100 (block 1202). More specifically, theturbo-compressor 230 is coupled to the engine 200 via the dual-inletbleed air system 1002. In particular, the compressor inlet 236 of theturbo-compressor 230 is fluidly coupled to a low-pressure bleed airsource via, for example the low-pressure port 238 and the first inletpassageway 1008 (block 1204). The compressor inlet 236 of theturbo-compressor 230 is fluidly coupled to an intermediate pressurebleed air source via, for example, the intermediate port 1012 and secondinlet passageway 1010 (block 1206). The turbine inlet 262 of theturbo-compressor 230 is fluidly coupled to the high-pressure bleed airsource of the engine 200 via, for example, the high-pressure port 266and the third inlet passageway 1014 (block 1208), and the turbine inlet262 of the turbo-compressor 230 is fluidly coupled to an intermediatepressure bleed air source via, for example, the intermediate port 1012and the fourth inlet passageway 1016 (block 1210). Additionally, thefirst, second, third and fourth inlet passageways 1008, 1010, 1014, and1016 are fluidly isolated via one or more fluid control devices such as,for example, the flow control members 244, 1018, 268 and 1020. (block1212).

In some examples, the compressor inlet 236 and the turbine inlet 262 maybe fluidly coupled to a common intermediate pressure bleed air source.In other examples, the compressor inlet 236 and the turbine inlet 262may be fluidly coupled to different intermediate pressure bleed airsources.

Although certain example methods, apparatus and articles of manufacturehave been described herein, the scope of coverage of this patent is notlimited thereto. On the contrary, this patent covers all methods,apparatus and articles of manufacture fairly falling within the scope ofthe appended claims either literally or under the doctrine ofequivalents.

1. An apparatus comprising: a turbo-compressor including: a compressorhaving a compressor inlet fluidly coupled to (1) a bleed air port of alow-pressure compressor of an aircraft engine and (2) an intermediateport of a high-pressure compressor of the aircraft engine, thecompressor inlet to receive fluid from either the low-pressurecompressor or the high-pressure compressor based on a first systemparameter of the aircraft; and a turbine having a turbine inlet fluidlycoupled to (1) the intermediate port of the high pressure compressor ofthe aircraft engine and (2) a high-pressure port of the high pressurecompressor of the aircraft engine, the turbine inlet to receive fluidfrom either the intermediate port of the high-pressure compressor or thehigh-pressure port of the high-pressure compressor based on a secondsystem parameter of the aircraft, the compressor and the turbine of theturbo-compressor being separate from the aircraft engine.
 2. Anapparatus of claim 1, wherein the first system parameter issubstantially similar to the second system parameter.
 3. An apparatus ofclaim 2, wherein each of the first and second system parameterscomprises a speed of the aircraft.
 4. An apparatus of claim 1, whereinthe first system parameter comprises a first pressure value and thesecond system parameter comprises a second pressure value.
 5. Anapparatus of claim 4, wherein the compressor inlet is to receive fluidfrom the low-pressure compressor when a pressure of the fluid from thelow-pressure compressor is greater than the first pressure value and thecompressor inlet is to receive fluid from the intermediate port of thehigh-pressure compressor when a pressure of the fluid from thelow-pressure compressor is less than the first pressure value.
 6. Anapparatus of claim 4, wherein the turbine inlet is to receive fluid fromthe intermediate port of the high-pressure compressor when a pressure ofthe fluid from the high-pressure port is greater than the secondpressure value and the turbine inlet is to receive fluid from thehigh-pressure port of the high pressure compressor when a pressure ofthe fluid from the high-pressure port is less than the second pressurevalue.
 7. An apparatus as defined in claim 1, further comprising a firstinlet passageway to fluidly couple the low-pressure port and thecompressor inlet and a second inlet passageway to fluidly couple theintermediate port and the compressor inlet.
 8. An apparatus as definedin claim 7, further comprising a third inlet passageway to fluidlycouple the high-pressure port and the turbine inlet and a fourth inletpassageway to fluidly couple the intermediate port and the turbineinlet.
 9. An apparatus as defined in claim 8, further comprising flowcontrol members to control fluid flow between the first, second, thirdand fourth fluid inlet passageways.
 10. An apparatus as defined in claim9, wherein the flow control members comprise a first check valve fluidlycoupled to the first inlet passageway between the low-pressure port andthe compressor inlet, and a first shut-off valve fluidly coupled to thesecond inlet passageway between the intermediate port and the compressorinlet.
 11. An apparatus as defined in claim 10, wherein the flow controlmembers comprise a second shut-off valve fluidly coupled to the thirdinlet passageway between the high-pressure port and the turbine inlet,and a second check valve fluidly coupled to the fourth inlet passagewaybetween the intermediate port and the turbine inlet.
 12. An apparatuscomprising: a turbo-compressor comprising a compressor and a turbine,the compressor to increase a pressure of bleed air extracted from anaircraft engine at a first lower pressure to a second higher pressure toprovide pressurized bleed air to a system of the aircraft; a first inletpassageway to fluidly couple a low-pressure port from the aircraftengine to a compressor inlet of the compressor; a second inletpassageway to fluidly couple a first intermediate port from the aircraftengine to the compressor inlet; a third inlet passageway to fluidlycouple a high-pressure port from an aircraft engine to a turbine inletof the turbine; and a fourth inlet passageway to fluidly couple a secondintermediate port from the aircraft engine to the turbine inlet.
 13. Anapparatus as defined in claim 12, wherein the first intermediate portand the second intermediate port are common ports.
 14. An apparatus asdefined in claim 12, wherein fluid flow through the first and secondinlet passageways is controlled independently from the fluid flowthrough the third and fourth inlet passageways.
 15. An apparatus asdefined in claim 13, further comprising a first valve fluidly coupled tothe first inlet passageway between the low-pressure port and thecompressor inlet, a second valve fluidly coupled to the second inletpassageway between the intermediate port and the compressor inlet,wherein operation of the second valve between an open position and aclosed position causes the first valve to move between an open positionand a closed position.
 16. An apparatus as defined in claim 15, furthercomprising a third valve fluidly coupled to the third inlet passagewaybetween the high-pressure port and the turbine inlet, a fourth fluidvalve fluidly coupled to the fourth inlet passageway between theintermediate port and the turbine inlet, wherein operation of the thirdvalve between an open position and the closed position causes the fourthvalve to move between an open position and a closed position.
 17. Anapparatus as defined in claim 16, wherein the first inlet passageway andthe second inlet passageway converge upstream from the compressor inletand downstream from the first and second valves, and the third inletpassageway and the fourth inlet passageway converge upstream from theturbine inlet and downstream from the third and fourth valves.
 18. Amethod comprising: fluidly coupling a compressor inlet of aturbo-compressor to a low-pressure bleed air source provided by alow-pressure compressor of an aircraft engine via a first inletpassageway and fluidly coupling the compressor inlet to an intermediatebleed air source provided by a high-pressure compressor of the aircraftengine via a second inlet passageway; and fluidly coupling a turbineinlet of the turbo-compressor to a high-pressure bleed air sourceprovided by the high-pressure compressor of the aircraft engine via athird inlet passageway and fluidly coupling the turbine inlet and theintermediate bleed air source provided by high-pressure compressor via afourth inlet passageway, the aircraft engine comprising a housing thatcontains the low-pressure compressor and the high-pressure compressor ofthe aircraft engine, wherein the low-pressure bleed air source isprovided by a low-pressure bleed air port in the housing of the aircraftengine, the intermediate bleed air source is provided by an intermediatebleed air port in the housing of the aircraft engine, and thehigh-pressure bleed air source is provided by a high-pressure bleed airport in the housing of the aircraft engine.
 19. A method of claim 20,further comprising causing the compressor inlet to receive bleed airfrom the low-pressure bleed air source and the turbine inlet to receivebleed air from the intermediate bleed air source during a flightcondition of the aircraft, and causing the compressor inlet to receivebleed air from the intermediate bleed air source and the turbine inletto receive bleed air from the high-pressure bleed air source during anon-flight condition when the aircraft engine is idle.
 20. A method ofclaim 18, further comprising isolating the first, second, third andfourth inlet passageways via flow control devices.